The present disclosure relates to components for a gas turbine engine and, more particularly, to cooling features within an airfoil therefor.
Gas turbine engines typically include a compressor section to pressurize airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. Gas path components, such as turbine blades, often include airfoil cooling that may be accomplished by external film cooling, internal air impingement and forced convection either separately or in combination. In forced convection cooling, compressor bleed air flows through internal cavities of hot section blades and vanes to continuously remove thermal energy. Compressor bleed air enters the internal cavities through one or more inlets to the internal cavities, which then discharge though various exits. The internal cavities often communicate with a trailing edge cavity that directs cooling air around an internal pedestal array to axially exit through a trailing edge passage of the blade. Although effective, trailing edge tip corner burning/creep is common in turbine blades.
Advances in casting, such as refractory metal core (RMC) technology, facilitate significantly smaller and more complex passages to accommodate the elevated temperatures with a reduced flow of cooling air. Refractory metal cores are metal based casting cores usually composed of molybdenum with a protective coating. The refractory metal provides more ductility than conventional ceramic core materials while the coating (usually metallic) protects the base metal form alloying with the refractory metal in the investment casting process. RMCs have shown significant promise in casting feature sizes and geometries not attainable with ceramic cores alone.